The present invention relates generally to gas turbine engines, and, more specifically, to turbine rotor blades therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the combustion gases by turbine rotor blades which in turn power the compressor, and an upstream fan in an exemplary turbofan aircraft engine application.
Turbine blades typically increase in radial size in the downstream direction as energy is extracted from the combustion gases. And, high pressure turbine blades are typically hollow and provided with internal cooling circuits or channels through which a portion of the pressurized compressor air is channeled for cooling the blades during operation in the environment of the hot combustion gases.
Each rotor blade includes an airfoil extending radially outwardly from an inner platform, with the platform being joined by a shank to a supporting dovetail mounted in a corresponding slot in the perimeter in a supporting rotor disk. The flow channels extend through the airfoil, platform, and dovetail and typically include a plurality of inlets in the base of the dovetail for receiving the pressurized cooling air from the compressor.
During operation, the blades drive the rotor at substantial speed and are subject to centrifugal forces or loads which pull the blades radially outwardly in their supporting slots in the perimeter of the rotor disk. The dovetail typically includes multiple lobes or tangs that carry the centrifugal loads of each blade into the rotor disk while limiting the stresses in the blade for ensuring long blade life.
Each rotor blade is also subject to pressure and thermal loads and stresses from the combustion gases which flow thereover during operation. And, the blades are also subject to vibratory stress due to the dynamic excitation thereof by the rotating blades and the pressure forces from the combustion gases.
Since the turbine airfoil is relatively thin for minimizing weight and resultant centrifugal loads, it is subject to the vibratory excitation in various modes. For example, the airfoil is subject to vibratory bending along the radial or longitudinal span thereof, as well as higher order bending modes along the axial chord direction.
Accordingly, turbine blades may also include a suitable vibration damper suitably mounted under the blade platforms. The dampers are supported by the platform and dovetail and add centrifugal loads to the rotor disk. The dampers use friction with the excited platform to provide effective damping of the blade during operation at speed.
However, these dampers have limited effectiveness for the various modes of vibration of the turbine blade during operation, including the higher order natural modes of airfoil vibration which involve complex combinations of airfoil bending in both the chord and span directions.
Accordingly, it is desired to provide an improved damper for a gas turbine engine turbine blade.